Gas turbine engine, nacelle thereof, and associated method of operating a gas turbine engine

ABSTRACT

The nacelle can have an inlet fluidly connecting a main gas path of a gas turbine engine core, the inlet having an inlet edge connecting an external skin to an internal duct wall, and a step formed in a surface of at least one of the skin and the duct wall, the step delimiting a first portion of the surface from a second portion of the surface, the second portion of the surface being recessed relative to the first portion of the surface, the second portion of the surface extending away from both the step and the inlet edge, whereas the first portion of the surface extends between the inlet edge and the step.

TECHNICAL FIELD

The application relates generally to gas turbine engines and, moreparticularly, to ice mitigation systems therefor.

BACKGROUND OF THE ART

This engine nacelle skins, which are exposed to the environment, may besubject to ice accumulation. In the case of engine nacelles, iceaccumulation in the vicinity of the inlet can be particularlyundesirable as accumulating ice can eventually separate from the surfaceand represent a source of foreign object damage (FOD). To mitigate iceaccumulation to the inlet portion of engine nacelles, it was known toprovide heating within the nacelle, such as via hotter air bled from thecompressor for instance. Although known systems were satisfactory to acertain extent, there always remains room for improvement, such as inreducing the amount of heat required to achieve the intended purpose forinstance.

SUMMARY

In one aspect, there is provided an aircraft engine nacelle comprisingan inlet fluidly connecting a main gas path of a gas turbine enginecore, the inlet having an inlet edge connecting an external skin to aninternal duct wall, and a step formed in a surface of at least one ofthe skin and the duct wall, the step delimiting a first portion of thesurface from a second portion of the surface, the second portion of thesurface being recessed relative to the first portion of the surface, thesecond portion of the surface extending away from both the step and theinlet edge, whereas the first portion of the surface extends between theinlet edge and the step.

In another aspect, there is provided a method of operating a gas turbineengine, the method including a flow of air circulating along a surfaceof an inlet portion of the gas turbine engine, the flow of air drawingwater droplets along the surface until the water droplets reach an edgeof a step leading to a recessed portion of the surface, the flow of airseparating the water droplets from the surface at the edge of the step.

In another aspect, there is provided an aircraft engine nacellecomprising an inlet fluidly connecting a main gas path of a gas turbineengine core, the inlet having an inlet edge connecting an external skinto an internal duct wall, and means for a flow of air circulating alongthe internal duct wall to draw water droplets along the surface untilthe water droplets reach an edge, and to separate the water dropletsfrom the surface at the edge of the step.

In another aspect, there is provided an aircraft engine nacellecomprising an inlet fluidly connecting a main gas path of a gas turbineengine core, the inlet having an inlet edge connecting an external skinto an internal duct wall, and means for a flow of air circulating alongthe external skin to draw water droplets along the surface until thewater droplets reach an edge, and to separate the water droplets fromthe surface at the edge of the step.

In a further aspect, there is provided an aircraft engine comprising agas turbine engine core having a main gas path extending, in serial flowcommunication, across a compressor section, a combustor, and a turbinesection, the gas turbine engine core housed within a nacelle, thenacelle having an inlet fluidly connecting the main gas path, the inlethaving an inlet edge connecting an external skin to an internal ductwall, and a step formed in a surface of at least one of the skin and theduct wall, the step delimiting a first portion of the surface from asecond portion of the surface, the second portion of the surface beingrecessed relative to the first portion of the surface, the secondportion of the surface extending away from both the step and the inletedge.

DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a schematic cross-sectional view of a gas turbine engine;

FIG. 2A shows an inlet portion of a nacelle of the gas turbine engine ofFIG. 1;

FIG. 2B is a portion 2B-2B of FIG. 2A, shown enlarged;

FIG. 2C is a portion 2C-2C of FIG. 2A, shown enlarged; and

FIG. 3 is a heating system of the engine of FIG. 1.

DETAILED DESCRIPTION

FIG. 1 illustrates a gas turbine engine 10 of a type preferably providedfor use in subsonic flight, generally comprising in serial flowcommunication an inlet 20, a fan 12 through which ambient air ispropelled, a compressor section 14 for pressurizing the air, a combustor16 in which the compressed air is mixed with fuel and ignited forgenerating an annular stream of hot combustion gases, and a turbinesection 18 for extracting energy from the combustion gases, with rotarycomponents rotating around a main axis 11. The gas turbine engine 10 ishoused in a nacelle 22, which has an aerodynamically shaped externalsurface. In this example, the nacelle 22 forms an enclosure which isdistinct from the passenger compartment of the aircraft, and morespecifically, the nacelle 22 is separated from the passenger compartmentby a portion of a wing of the aircraft (not shown).

Gas turbine engines 10 typically have a main gas path extending throughthe compressor section 14, combustor 16 and turbine section 18. Turbofanengines, in particular, have a bypass path formed around the coreengine, within the nacelle 22. In any case, the inlet fluidlycommunicates with the main gas path, and in this case, it alsocommunicates with the bypass path. The shape of the nacelle 22 dependson the type of engine and is typically selected in a manner toaccommodate the specifics of the engine.

An inlet portion 20 of an example nacelle 22 is shown enlarged in FIG.2A. In this example, the inlet portion 20 has an inlet edge 102 which,during flight, separates a portion of the flow which is directed intothe engine 10, a sub-portion of which will flow along an internal ductwall 104 of the engine 10, from a portion of the flow which is directedaround the nacelle 22, a sub-portion of which will flow along anexternal skin 106 of the nacelle 22. In this embodiment, the inlet edgeis rounded.

The example inlet portion 20 is provided with a heater internal to therounded portion 108. The heater can be a heating air conduit 110 havinga plurality of apertures dissipating hotter air bled from thecompressor, for instance, or any suitable heater. The heater can be usedto heat water, in solid or liquid phase, which comes into contact withthe inlet portion 20, to avoid it forming and accumulating a layer ofice, which could eventually dislodge and represent a potential FOD. Thepower directed to the heater can be modulated as a function of theamount of power expected to be required to achieve this purpose, forinstance.

However, the heater has a limited range, and even if, for a given power,it can avoid ice accumulation within its range, liquid water circulatingalong the surface can eventually exit its range and form an iceaccumulation downstream of its range. To avoid this, the range of theheater can be extended, to a certain extent, by supplying additionalpower (hotter water will travel farther before freezing, especially ifit runs along a warmer surface), but this is done at the cost of theadditional power, which is typically undesired. Moreover, someembodiments may have practical limitations to the amount of extension ofheater range achievable by added power.

FIG. 2B presents an example embodiment where a step 112A is provided inthe surface 115 along which the liquid water circulates, in a mannerthat as the liquid water droplets 113 reach the edge of the step 112A,its velocity, entrained by the air velocity and viscosity, entrains itsseparation, and ejection, from the surface 115, after which it canremain entrained in the air flow rather than freezing and accumulatingonto a cooler portion of the surface, to eventually detach and causeFOD. Indeed, small droplets of water, even when solidified into smallice fragments, can have insufficient mass to cause any damage to theengine, by contrast with larger ice accumulations.

More specifically, in the example presented in FIG. 2B, the step 112A isformed in the duct wall 104 of the gas turbine engine 10, in thevicinity of the inlet edge 102. The step 112A can be said to form adiscontinuity in the surface 115, or to more specifically delimit arecessed (or second) portion 114 of the surface 115 from a non-recessed(or first) portion 116 of the surface 115. The recessed portion 114 ofthe surface 115 is offset, at the step 112A, from the non-recessedportion 116 of the surface 115 by a distance equivalent to the “height”of the step 112A. The recessed portion 114 is recessed relative to theair flow. The step 112A faces downstream relative to the movement of thewater along the surface, in the sense that if an imaginary Lilliputianperson would walk and go up the step, he would be walking against thewind flow, whereas if he would walk and go down the step, he would havethe wind in its back. Otherwise said, the recessed portion 114 extendsfrom the step 112A both away from the step 112A and the rounded portion102, whereas the non-recessed portion 116 extends between the roundedportion 102 and the step 112A.

The height of the step 112A can vary greatly depending on the size ofthe engine and the specifics of the embodiment. However, for the purposeof providing an order of magnitude, it can be said here that the heightof the step 112A can be expected to be between 0.010″ and 0.200″ in mostpractical applications. Greater heights may represent a flow distortionjudged as being too large, while not providing sufficient compensatingadvantages, whereas a height smaller than 0.010″ may not be sufficientto cause ejection of the water droplets 113. The exact height for aspecific application can be determined based on simulation or testing,for instance. Similarly, the sharpness of the step, i.e. the dimensionof the fillet radius of the edge of the step, can vary greatly from oneembodiment to another and can be chosen in view of optimizing theefficiency of a specific embodiment. Typically, the ratio of the filletradius to the height of the step can be between 0 and 1, and the filletradius can thus be less than 0.200″, for example.

In the specific embodiment illustrated in FIG. 2B, the step 112A has ariser in the form of a riser portion of the surface, which extendsnormal to the recessed portion 114 of the surface 115, along a distancecorresponding to the height of the step 112A. The riser faces downstreamrelative to the movement of the water droplets 113, or in this specificembodiment, rearwardly relative to the orientation of thrust of theturbofan engine.

In the case of a turbofan engine, providing a step 112A along the bypassduct wall, as opposed to along the nacelle skin, can be particularlyuseful in avoiding ice accumulation which would be likely to otherwiseform a potential source of FOD.

Indeed, ice accumulating on the nacelle skin 106 will typically notrepresent a potential source of FOD during flight, however it can stillbe undesired for other reasons. It will be noted that the flow dynamicsduring takeoff are very different than during flight, hence if ice hasaccumulated on a nacelle skin 106 on a previous flight and remains inthe vicinity of the inlet, a sufficient velocity of air may be drawnforwardly along the skin, towards the inlet, during the next takeoff,causing detachment of the ice accumulation and a FOD. Accordingly,several reasons may motivate the use of a backward facing step 112B onthe external skin 106 of the nacelle 22 in addition to, or perhaps eveninstead of, a step 112A on the internal duct wall 104. This backwardfacing step 112B on the external skin 106 can be used to eject waterrunning along the surface during flight, prevent the ejected water fromfreezing and forming an ice accumulation during flight, and thus preventeventual aspiration of such an ice accumulation by a reverse flowoccurring during takeoff, for instance.

FIG. 2C shows an example of a nacelle inlet 20 having a step 112B formedon the external skin 106 on the nacelle 22.

It will be understood that in the specific case of a turbofan gasturbine engine, the inlet 20 extends annularly around the engine's mainaxis 11, and therefore the inlet edge 102, skin 106, and duct wall 104can be axisymmetric around the main axis 11. In such a context, the stepcan be designed in a manner to extend around the entire circumference ofthe inlet 20, for instance. However, in some embodiments, it may bedetermined that one or more targeted circumferential portions of theinlet 20 are more prone to ice accumulation, and the step can bedesigned to extend only partially around the circumference, incoincidence with the one or more circumferential portions more prone toice accumulation. In the case of a turbofan gas turbine engine, the ductwall 104 can be a an outer bypass duct wall for instance. It will benoted, however, that in alternate embodiments, the step can be providedon nacelle inlets 20 of gas turbine engines 10 having other geometries,such as different types of gas turbine engines 10, and the step can thusbe adapted accordingly.

Returning to the illustrated example of a turbofan gas turbine engineapplication, the inlet edge 102 can form part of a D-duct 108 connectedto a remainder of the nacelle 22, or bypass duct, as known in the art,and the heating conduit 110 can extend circumferentially within theD-duct 108, for instance. In such an embodiment, the step can coincidewith, and be formed by, the junction between the D-duct 108 and adjacentsections of the nacelle 22, for instance. The heating conduit 110 canhave a plurality of apertures forming heating air outlets, and beconnected to a compressor to receive bleed air therefrom. An example ofa possible arrangement is shown in FIG. 3, where the annular heatingconduit 210 is shown to be connected, via a thermally insulated pipesegment 220, to an engine bleed port 222, and such an arrangement canhave a pressure regulating and shut-off valve (PRSOV) 224 associatedwith the thermally insulated pipe segment 220, for instance.

Returning to FIGS. 2A and 2B, and the specific context of a turbofanengine, it is common for turbofan engines to have outer bypass ductsintegrating acoustic panels 118 in a manner to impede sound transmissionfrom the core engine to the passengers. In such a scenario, the step112A can be located between the inlet edge 102, and the acoustic panel118, for instance.

The above description is meant to be exemplary only, and one skilled inthe art will recognize that changes may be made to the embodimentsdescribed without departing from the scope of the invention disclosed.Still other modifications which fall within the scope of the presentinvention will be apparent to those skilled in the art, in light of areview of this disclosure, and such modifications are intended to fallwithin the appended claims.

1. An aircraft engine nacelle comprising an inlet fluidly connecting toa main gas path of a gas turbine engine core, the inlet having an inletedge connecting an external skin to an internal duct wall, and a stepformed in a surface of at least one of the external skin and the ductwall, the step delimiting a first portion of the surface from a secondportion of the surface, the second portion of the surface being recessedrelative to the first portion of the surface by a height of the step,the second portion of the surface extending away from both the step andthe inlet edge, whereas the first portion of the surface extends betweenthe inlet edge and the step.
 2. The aircraft engine nacelle of claim 1wherein the height of the step is of between 0.010″ and 0.200″ measurednormal to the surface.
 3. The aircraft engine nacelle of claim 1 whereinthe step has a riser.
 4. The aircraft engine nacelle of claim 1 whereinthe aircraft engine is a turbofan engine, and the duct wall is an outerbypass duct wall.
 5. The aircraft engine nacelle of claim 4 wherein theinlet edge is a portion of a D-duct, the D-duct connecting the skin andthe duct wall.
 6. The aircraft engine nacelle of claim 5 wherein thestep is formed at a junction between the D-duct and the duct wall. 7.The aircraft engine nacelle of claim 5 wherein a heating air conduit isprovided inside the D-duct, the heating air conduit having a pluralityof heating air outlets, and being connected to a compressor bleed airsource.
 8. The aircraft engine nacelle of claim 4 wherein the inner ductwall has an acoustic panel, the step being located along the surface,between the inlet edge and the acoustic panel.
 9. The aircraft enginenacelle of claim 1 wherein the step is formed in the duct wall.
 10. Theaircraft engine nacelle of claim 8 further comprising an other stepformed in the skin, the other step delimiting a recessed portion of theskin, the recessed portion of the skin extending away from both theother step and the inlet edge.
 11. The aircraft engine nacelle of claim1 wherein the inlet edge, skin and duct are annular.
 12. The aircraftengine nacelle of claim 11 wherein the step is backward facing.
 13. Amethod of operating a gas turbine engine, the method including a flow ofair circulating along a surface of an inlet portion of the gas turbineengine, the flow of air drawing water droplets along the surface untilthe water droplets reach an edge of a step leading to a recessed portionof the surface, the flow of air separating the water droplets from thesurface at the edge of the step.
 14. The method of claim 13 wherein themethod further comprises subjecting water in solid state to heating, andthereby transforming the water in solid state into the water droplets.15. The method of claim 13 further comprising directing the detachedwater droplets into one of an engine core main gas path, or a bypassduct.
 16. An aircraft engine comprising a gas turbine engine core havinga main gas path extending, in serial flow communication, across acompressor section, a combustor, and a turbine section, the gas turbineengine core housed within a nacelle, the nacelle having an inlet fluidlyconnecting the main gas path, the inlet having an inlet edge connectingan external skin to an internal duct wall, and a step formed in asurface of at least one of the skin and the duct wall, the stepdelimiting a first portion of the surface from a second portion of thesurface, the second portion of the surface being recessed relative tothe first portion of the surface, the second portion of the surfaceextending away from both the step and the inlet edge.
 17. The aircraftengine of claim 16 wherein the step has a height of between 0.010″ and0.200″ measured normal to the surface.
 18. The aircraft engine of claim16 wherein the aircraft engine is a turbofan engine, and the duct wallis an outer bypass duct wall.
 19. The aircraft engine of claim 16wherein the step is formed in the duct wall.
 20. The aircraft engine ofclaim 16 wherein the inlet edge, skin and duct are annular.